Rotary machines such as gas turbine engines include a turbine section. The turbine section has a rotor assembly which includes a rotor disk and a plurality of rotor blades extending radially outwardly from the disk. A flowpath for hot working medium gases extends axially through the rotor assembly and between the rotor blades of the rotor assembly. In modern gas turbine engines the temperature of the working medium gases may approach twenty-five hundred degrees fahrenheit (2,500.degree. F.) at the entrance to the turbine.
Each rotor blade in the forward portion of the turbine has an airfoil section which extends radially outwardly from the rotor assembly and into the working medium flowpath. The airfoil is defined by a plurality of airfoil sections disposed about a spanwisely extending axis. The airfoil adapts the rotor blade to extract energy from the hot working medium gases for driving the rotor assembly about an axis of rotation. Accordingly, each rotor blade is bathed in the hot working medium gases and receives heat from these gases. Heating of the rotor blade causes thermal stresses in the airfoil which, when combined with mechanical stresses resulting from rotation of the blade about the axis of rotation decrease the structural integrity and fatigue life of the airfoil.
Rotor blades at the forward portion of the turbine are coolable to preserve the structural integrity and fatigue life of the airfoil. Cooling air is flowed through passages on the interior of the rotor blade and to the airfoil to remove heat from the airfoil. The cooling air is discharged at the downstream end of the airfoil through openings and from film cooling holes which extend through the surface of the airfoil at preselected locations on the airfoil surface. The cooling air provides convective cooling to the airfoil wall as the cooling air is flowed through the openings. These cooling air holes, commonly referred to as "film cooling holes" also discharge cooling air at the surface of the airfoil. The cooling air extends as a film over a portion of the airfoil at critical locations to provide a layer of cool air which blocks the hot working medium gases from contacting the surface of the airfoil, thus decreasing the transfer of heat to the airfoil at critical locations.
One example of an airfoil in which cooling air is flowed through the interior of the airfoil and discharged to the exterior at critical locations is shown in U.S. Pat. No. 4,474,532 issued to Pazder, entitled Coolable Airfoil for a Rotary Machine. In Pazder, the airfoil includes a leading edge 34 and a trailing edge 36, a suction sidewall 38, and a pressure sidewall 42 are joined at the leading edge region and the trailing edge region. In the leading edge region, cooling air holes extend from the interior to the exterior to discharge cooling air in this region of the airfoil.
Another example is shown in U.S. Pat. No. 4,753,575 issued to Levengood et al., entitled Airfoil with Nested Cooling Channels. As shown in Levengood, the forward most portion of the airfoil may have three (3) or more cooling air holes as illustrated in FIG. 2 and FIG. 6 which correspond to the sections 2--2 (three (3) cooling air holes) in FIG. 1 and the section 6-6 (four (4) cooling air holes) in FIG. 5.
In modern gas turbine engines, the film cooling holes in the leading edge region are typically located relative to the predicted stagnation point of gas path air. The stagnation point for any airfoil section under a particular operative condition is the point on the airfoil at which the hot working medium gases impact the airfoil and have the minimum (theoretically zero) velocity. Working medium gases impacting the airfoil on one chordwise side of the stagnation point are flowed toward the trailing edge of the airfoil on the suction side of the airfoil. Working medium gases impacting the airfoil on the other chordwise side of the stagnation point flow are flowed toward the trailing edge on the pressure side of the airfoil. A line connecting the series of stagnation points in the spanwise direction forms the aerodynamic leading edge of the airfoil. As will be realized, the aerodynamic stagnation point is a function of the angle of attack of the approaching flow. The angle of attack changes for different operative conditions of the engine and varies the location of the aerodynamic leading edge. This is different from the forward most portion of an airfoil section which is commonly referred to as the mechanical stagnation point of the airfoil section. A line connecting the mechanical stagnation points of the airfoil forms the mechanical leading edge (commonly referred to simply as the "leading edge") of the airfoil. The location of this leading edge never changes. Because modern airfoils encounter pressure and temperature gradients in the oncoming flow and because the airfoils do have a leading edge which is angled somewhat in a spanwise direction to a radial line, the aerodynamic leading edge tends to curl around the leading edge in an S shape.
The prior art teaches disposing rows of film cooling holes in the airfoil in the leading edge region in a way that avoids the line of the most troublesome aerodynamic stagnation points. This aerodynamic leading edge connecting these points occurs at the operative condition of the turbine that has the highest temperature in the gas path (sea level take off, hot day). In the prior art, the rows of leading edge holes are typically disposed parallel to the location of this aerodynamic leading edge and each row is spaced chordwisely an equal distance from the aerodynamic leading edge. This avoids having film cooling holes at a point where the greatest amount of dynamic velocity pressure is converted to static pressure. If this were to occur on the row of film cooling holes, the high static pressure might block the cooling air from flowing outwardly in the film cooling hole.
During the operative life of the engine, the aerodynamics of the turbine may change because of stator vane or rotor blade re-staggering to accommodate growth variations of the engine or other aerodynamic concerns. It is a simple matter, in such cases, to change slightly the angle of the root of the rotor blade with respect to the axis of the engine, changing the angle of attack of the airfoil with respect to the oncoming flow. This may shift the aerodynamic leading edge chordwisely so that it is aligned with a row of film cooling holes and blocks the flow of cooling air. The blockage requires a redesign of the cooling scheme to shift the cooling air holes away from the aerodynamic leading edge. A considerable amount of work in design and manufacturing is required to carry out this change. The task is even more complex for airfoils having internal impingement cooling of the leading edge region because the internal cooling requirements impose constraints on the location of holes extending through the airfoil.
Accordingly, scientists and engineers working under the direction of Applicant's assignee have sought to develop new cooling schemes for the leading edge region of the airfoil which are better able to accommodate changes in the location of the aerodynamic leading edge that result from an aerodynamic redesign of the turbine or a change in operative conditions of the engine.